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How did the controls work for variable-pitch props?
Variable-pitch propellers date way back to the 1930's and even earlier. I would like to know, what kind of controls they used? Some kind of lever or crank? How heavy were these controls? I'm thinking of the drag the propeller has, since even near landing, the prop could be spinning very fast and would be resistant to changing it's pitch.
I'm interested in the old days, before modern electrical control by wire.
Note: I am not asking how the variable-pitch mechanisms work. I might ask that in a separate question.

Related. You may include your current state of research as explained in the help center.

Why would the propeller be resistant to changing pitch? The blades rotate around axis through their aerodynamic center and the centrifugal force is parallel to that axis, so it does not affect rotation around it either. The only force opposing the control is the feathering spring or counterweight that feathers the propeller in case of failure of the mechanism.

Note that how the mechanism works determines how the controls may work. So while you may not be asking it, it is still very relevant.

@JanHudec A feathered propeller has the least drag in a glide, while a normally-pitched prop has the most drag. So there must be an increase in drag if you're changing the blade-pitch to be more "steep". That increase in drag would resist the pitch change just like control surfaces are resistant to deflection.

Control by wire? As in electrically?
Many of the small planes like Cessna Cardinal use an engine like Lycoming O-360 (4 cylinder) use a spring to return the blades to a flat pitch, and engine oil pressure to change the blade pitch. The control knob in the cockpit mechanically allows the oil pressure to change the pitch.  In other small planes, like a Bonanza, there is a set of counterweights instead of a spring to accomplish the same. No electric control at all.

Can you tell me how big is this control knob and how many turns is needed for a change of x angles of blade pitch?

In Cessnas, it is a knob, maybe 1.5inch in diamater.  https://www.aircraftspruce.com/pages/ap/cessna_controls/cessnapropcontrols2.php   I unwind it (or pull it out) to increase the pitch until I see 2400 RPM in cruise. Full forward for takeoff.  On descent for landing, I work both the RPM and manifold pressure, when finally slow enough on speed the RPMs are driven by the throttle directly and pitch control has no effect, and then the blue knob goes all the way in. I can't say on the number of turns and resulting pitch change.


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Has there ever been an uncommanded lowering of landing gear?
Has there ever been an uncommanded lowering of the landing gear on a large commercial jet? I could not find any instance of it in a preliminary google search. Could a simple short-circuit cause such a thing?
I think this would actually be pretty bad at cruising conditions, because the airplane is going much faster up there, than on approach. But the air is also much thinner so I can't be sure.

IIRC (so take this with a large handful of salt), jetliners cruise at something like 300-400 KIAS, and land at maybe 100-120 KIAS. Since it's indicated airspeed that would be of interest here, it stands to reason that lowering of the landing gear would still be much worse in cruise than near landing.

@MichaelKjörling, I am rather confident jetliners cruise below 300 KIAS. The Vmo is usually 320 or 330 knots indicated, but the Mmo becomes the limit above around FL270 and pushes it below 300 knots indicated.

@JanHudec Yes, very much so. High 30's for a typical cruise altitude, Mach .78 ballpark for a 737, you could have 450 knots +/- TRUE airspeed, but 250ish knots INDICATED.  Dropping the gear there would be LOUD, but within its operating limits. With millions of commercial airline flights per year, the answer to "has there ever" is surely "yes". Noteworthy enough to get its own internet article, doubtful.

@JanHudec It's certainly possible. Like I said, apply a large handful of salt.

I have never heard of uncommanded gear extension/retraction. My guess is no it has never happened but a very thorough database search would be needed to confirm/disconfirm that.

Related

@kevin: Actually, it has... during a high-G pullout from an inverted supersonic dive.

The answer to your question is, in fact, "yes" - but the circumstances involved demonstrate just how difficult it is to make an aircraft's landing gear deploy without being commanded to do so.
In 1985, CI006, a 747SP, experienced an uncommanded flameout of its #4 (right outboard) engine (something that particular engine was quite prone to doing) while cruising at FL410 en route from TPE to LAX.  An attempted relight failed, and a lack of corrective flight control inputs combined with a poorly-designed autopilot caused the airplane to decelerate and gradually roll into a steep bank despite the autopilot's attempts to counter the thrust asymmetry; the captain eventually noticed this and disconnected the autopilot, but made no flight control inputs of his own.
This, unsurprisingly, resulted in an immediate loss of control, and the aircraft rolled into a steep inverted dive, during which it exceeded mach 1 for an indeterminate period of time, plunging 30 kilofeet before the captain was able to orient himself when the aircraft broke out of the clouds at 11 kft and pull out of the dive, eventually levelling the aircraft at 9,600 feet.  During the dive and subsequent pullout, the aircraft was subjected to aerodynamic loads exceeding +5 Gs,1 which, in addition to causing severe structural damage to the aircraft (including tearing away large parts of the aircraft's horizontal tail), ripped off the body gear doors, broke the support brackets holding the body gear uplock hooks in place, and forced the aircraft's left and right body gear into the down-and-locked position.  (The added drag from the extended body gear [and from the aircraft's 29-kft maximum gear-down altitude] forced the flight to divert to SFO, having insufficient fuel to reach LAX with almost half of the aircraft's landing gear hanging out in the airstream.)
So, yes, it has happened... during an over-5-G pullout from a 30-kft supersonic dive in a 747.

1: The highest normal acceleration recorded on the aircraft's FDR was +5.1 Gs during the pullout from the dive; however, the absolute maximum G-forces encountered by the aircraft are unknown, as the extreme G-forces exerted on the aircraft, and the FDR, caused the latter to fail to record properly for large portions of the dive and pullout.

“despite the autopilot's attempts to counter the thrust asymmetry”—the thing is that autopilot doesn't know how to do that. It controls only pitch and roll and yaw damper independently controls rudder to damp the dutch roll only, so it won't apply the rudder as is needed for outboard engine failure.

@JanHudec: In which case the autopilot was very poorly designed indeed, if it couldn't handle something as simple as asymmetric-thrust compensation.

as far as I know all autopilots, by design and very intentionally, don't handle asymmetric thrust. It is something nobody wants to have in an autopilot. Airbus will even indicate how much rudder trim should be added, but it requires pilots to do it. It would also have to be a special function, because autopilot does not normally have any reason to manipulate rudder, and nobody wants that either.

@JanHudec: The A300/A310's autopilot does use the rudder to compensate for engine-failure-generated asymmetric thrust (see the footnote on page 19/33 of this AAR), and I would be surprised if they had removed that autopilot feature in their later airliners (given Airbus's automation fetish).

there is similar note in A320 materials that says it will indicate the correct amount of rudder trim, but won't add it. Which does not mean it does not compensate anything—the normal law will keep wings level, and the yaw damper will probably add some rudder, but the plane will be in slight skidding turn. Regarding deleting features, keep in mind that the A320 control system was completely designed from the ground to be consistent and simple. A lot of things are completely different from A300.

"because autopilot does not normally have any reason to manipulate rudder" - Ummm, autolanding?  (Both to maintain runway alignment during approach in a crosswind, and to maintain directional control during rollout.)

Rollout, yes. The ground mode is rather special. On approach airliners normally crab all the way to touch-down and autoland usually has cross-wind limit of just 10 knots, so it's unlikely to differ from the normal yaw damper logic.

That's pretty impressive. Is this the max speed achieved by a 747 that made it to a runway?

@OrganicMarble: Don't know, but I wouldn't be surprised.

Only thing I can find is on the X-15 rocket plane. Expansion of the airframe caused the nose gear to come unhooked and deploy at Mach 4.2. The air friction overheated the tires, damaging them and causing them to disintegrate on touchdown. The pilot was still able to make a safe landing.

"The air friction tires and caused them to disintegrate on touchdown." – Are you missing a verb near the beginning of that sentence?

@TerranSwett I’m missing something, but I don’t remember what I was trying to say.

@TomMcW: How's it look now?

Question asks specifically for "a large commercial jet"

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In the early days of flight, were there any cockpit control schemes other than the modern one?
Back in the early days, what kind of control schemes were invented for aircraft, other than the modern one? (pedals to yaw, yoke to pitch and roll).
I'm only interested in control schemes that can be done from the cockpit and control relatively small surfaces. So no weight-shifting, no wing-warping.
Hypothetical examples: 2 yokes (1 for each aileron and elevator), a steering wheel?, one-axis yokes?

Are you only interested in flight controls?

@Pilothead Flight controls and the surfaces they govern.

Note that quite a few modern planes don't have yokes.  They use joysticks https://en.wikipedia.org/wiki/Joystick#Aviation or side-stick controllers https://en.wikipedia.org/wiki/Side-stick

One of the EAA monthly videos showed the restoration of a very early airplane where the stick was pushed forward and aft to go right and left. Pushing was left and pulling was right!

Interesting question. I remember seeing a documentary (or maybe even a Top Gear segment) about really early cars and, while they converged surprisingly quickly to the modern control methods and layout, there were all kinds of weird and wonderful systems in the early days.

The Dunne D.8 used a pair of levers, each controlling one elevon.

Early stealth bomber?

@Harper certainly, not one was ever even acquired - let alone shot down - by any radar guided missile system.

@GrimmTheOpiner Why limiting to missiles? It has an unblemished war record: Zero machines were shot down, this is quite an achievement!

The early wright glider (1902) had its rudder connected to the wing warping system.

they connected the rudder control cables to the wing-warping hip
  cradle, so a single motion by the pilot operated both controls. They
  also changed the original double rudder to a single rudder, as shown
  here.

The wrights also used a hip cradle to move the wing warping system (early aileron system) as well as the cross connected rudder which I cant say exists any more.

The Wrights steered their 1903 Flyer with a hip cradle, which the
  pilot controlled by sliding his hips from side to side.

Similarly the elevator had a dedicated independent control.

This lever, operated by the pilot's left hand, was connected by a
  pulley system to the plane's elevator.

Early aircraft that sported a rotary engine, the style where the entire engine spins and the crank remains stationary generally had no effective throttle. Instead the throttle was controlled with a "blip switch" that grounded the magneto temporarily and cutting the engine out. This is far different than our modern throttle systems.
See the answers to this question r.e. steering wheels

Wait...  Didn't OP specifically state not to answer with wing-warping or weight shifting?

@Keeta the hip system also controlled the rudder which the OP stated was in scope.

So this one control does control both a small surface and a large surface.  To me this disqualifies the control since you could not control only the small surface with any method of the manipulation of the control.

The Wrights themselves experimented with a variety of frankly scary control arrangements after their elevator-lever-and-hip-cradle arrangement in the original Flyer.
The Flyer III had three levers: the lever on the pilot's outside right turned the rudder, the one to the inside right warped the wings, and the lever on the left controlled the elevator; the first Model A was also so equipped  [1]. Wilbur found this difficult, blaming a crash in the Flyer III on it, and for the Model A they took to France he used two levers: the left hand controlled the elevator, while the right hand controlled wing-warping with a fore-and-aft motion, and the rudder with a lateral movement.
For a while, you could order a Model A with either 'Orville' or 'Wilbur' controls, though the 'Orville' system was modified to more resemble the 'Wilbur' system by creating a jointed lever, controlling the wing-warping by moving the whole assembly fore-and-aft, and rudder by moving the upper section laterally. This stick was placed between the seats, while an elevator lever was put to the left of the left seat and the right of the right seat - so if you switched seats, the handedness of the controls also switched (to be fair, that is also the case today with center-mounted throttles.)
There were undoubtedly other variants wherever experimental aircraft were built. In 1909, Baron Pierre de Caters was using a foot-operated rudder bar, and the invention of the conventional stick-and-rudder arrangement is generally credited to Esnault-Pelterie and first used by Bleriot.
Then there are helicopters: collective pitch, cyclic pitch, torque pedals, and throttle (this last often twist-grip).
Various near-space aircraft (X-15, NF104A...) were fitted with thrusters for attitude control at altitudes where aerodynamic controls are ineffective. In the case of the NF104A (an F104 with a rocket motor added), control was through a separate 'stick' [5]:

Shaped like a pump handle of a yard spray tank, the RCS controller was
  rotated in the desired roll direction, pushed left and right for
  corresponding yaw and raised to nose up for increased alpha or down to
  lower it. Any and all command combinations were possible.

The rate of rotation around an axis was a function of how long the thruster had been activated, and continued after the control was centered, giving a very different feel than aerodynamic controls. It was not anticipated that the shut-down jet engine would continue to spin for some time in the very thin air, and would cause gyroscopic precession when the aircraft was rotated in pitch and yaw. To add to these difficulties, the thruster controller had been miswired for the first high-altitude flight, which nevertheless was flown successfully.

The Wright Flyer used a simple wooden lever in the left hand to control the forward mounted elevator for pitch control. A sliding body cradle was used to transmit left and right body movements to simultaneously control both the wing warping for roll control, and the rudder for yaw control.
This was the first 3 axis control.
With the Wright Flyer III the roll and yaw controls became 2 separate control inputs and the flight controls quickly evolved into the conventional stick and rudder that is still in use today.

Pre WW2 French Aircraft had the throttle reversed - i.e. Moving the throttle rearwards INCREASED engine power, and moving the throttle forwards DECREASED engine power. American Curtiss H-75 Hawks exported to France had the modifications made on re-assembly in France - something for any American company test-pilot in France to remember....

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In a pusher configuration with turboprop(s), how is the superhot exhaust avoided?
I actually looked around for turboprops in a pusher configuration, and so far I only found one: P180 Avanti. You can see the propellers are behind the turbine, and in fact it looks like two exhaust ports lead right into the propellers.
How is this possible? Turbine exhaust is very, very hot. "Superheated steam" is actually a technically accurate term. Won't this disturb the propellers and cause them to fail or else need huge maintenance? Is there some elegant way to avoid this geometry in a pusher configuration?

Beechcraft Starship also https://en.wikipedia.org/wiki/Beechcraft_Starship

Turbine exhaust was in fact a problem on one airplane, although not affecting the props.  The DeHavilland -7's overwing exhaust was found to be actually annealing the 7075 alloy upper wing planks and they were being found wavy in service from very slow compression buckling.  Dash 7s all have skin doublers behind the nacelles now.

Because the blade is moving in and out of the exhaust stream at high velocity, and the exhaust stream is perhaps 20% of the total exposure, the heat transfer to the blades is negligible from the standpoint of heating up aluminum enough to affect its heat treat, or heating epoxy to its transition temperature.  Meanwhile there are the anti-icing benefits.
There is an issue with the exhaust impinging on the prop on the Avanti, and there is a Hartzel Service Bulletin http://hartzellprop.com/wp-content/uploads/SB181A-R06-W.pdf that covers it.   The carbon and heat from the exhaust attacks the aluminum of the blade and the problem is blade corrosion once the paint finish starts to erode.  The SB is to inspect, clean up, and repaint the blades from this damage.  (Carbon and aluminum are at opposite ends of the galvanic scale and don't like to live together; this caused much grief on the CRJ program when someone who evidently missed their corrosion module in university decided to put carbon floor panels on aluminum support beams without an adequate barrier, with unfortunate results and driving a switch to titanium beams).
Really though, with the Avanti the biggest problem is the prop directly behind the gear acting like a FOD catcher for stuff thrown up by the tires.  You wouldn't want to try to fly an Avanti off a gravel strip, that's for sure.

Anti-icing is cool, but do you have a source for negligible heat transfer in terms of structural weakening? I would think it depends on how many blades and how thick they are. I've seen some really thick "fan" props, like on the Airbus A400M (although that's not pusher).

I'm going by the fact that the configuration was certified and in certification the margins of error are very large.  That being said, the potential is there and I'm starting to dimly remember some kind of issue with a pusher prop being damaged by exhaust heat, but can't recall the details.  I'll have to hunt around a bit.

Just remembered and edited my post.  Corrosion is the bugaboo, not annealing.

Any idea how long before corrosion becomes significant? I saw in the PDF it recommended 200 flight hours between inspections, but that's just inspections, not repairs. However, if repairs are just re-applying some paint, it doesn't sound that bad. BTW, what is CRJ program?

Canadair Regional Jet.  If you look at the initial issue of the Hartzel SB, March '93, it was only 3 years after the airplane was certified and there was likely a 6 month to a year investigation, so it probably started to show up within the first 2 years.  The main thing to control is paint erosion that exposes the bare aluminum to the carbon in the exhaust.  The location of the props behind the gear and right in the stream of grit particles thrown up, of which there are lots on paved runways, doesn't help.

For structural and corrosion problems, there is usually a 3x safety factor applied to test results.  Piaggio and Hartzel would have done some corrosion testing to simulate the in-service condition, and/or collected service data, and discovered that significant corrosion was happening in as little as 600 hours, so that would drive an inspection interval of 200.

I suppose one could route an exhaust stack from the engine installation to a location away from the engine or propellers.  But then again the benefits would have to outweigh the drawbacks of such a design.
Adding additional exhaust routing would solve the problems of structural damage due to heating, But it’s also bulkier, heavier and  imposes additional design requirements for heat shielding the aircraft structure, preventing exhaust gas leaks and designing an exhaust outlet installation to vent the exhaust gases away from the airframe somewhere else on the aircraft.  And that’s not really necessary if the temperature from the exhaust gases around the propeller blades is sufficiently cool enough to not cause structural damage the blades. A PT6 usually has a maximum Inter-turbine temperature (ITT) around 750° C but the  temperature of the exhaust gases exiting the exhaust stubs Is going to be lower-probably only order of 400°C or so. Add intermixing with outside air flow and the temperature of the exhaust gases may be down as low as 250 to 300°C by the time it passes through the propeller arc.   This is going to be far from hot enough to damage the propeller blades provided their manufactured out of the right material to the right specifications. It has an additional bonus and that the hot exhaust gases can serve as a means to prevent ice accumulation on  the propeller blades during flight, thereby forgoing a requirement for hot props or another kind of prop anti ice solution

It actually serves to heat the props and help in a de-icing situation so it appears to be quite deliberate and heat the props by design.

Propwash does not interfere with laminar flow over the wing and
  turbine exhaust gas heats the propeller blades, eliminating the need
  for de-icing.

However it does add to the exterior noise of the aircraft

The exterior noise level and its higher pitched sound has been exposed
  and presented to be the result primarily of the interaction of the
  turbine engine exhaust flows and the five-bladed pusher propellers.


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How much extra weight is added by strengthening a piston-prop fighter for carrier landings?
To land on a carrier, an aircraft must be strong enough to withstand sudden deceleration by the tail hook catching the arresting wire. I think it also requires strengthening the undercarriage to survive a harder "plop" on landing.
So I would like to know, how much extra weight is added due to both of these things?
I'm interested in the WW2-era piston props. The modern jet fighters are much heaver, so I'd rather save that for a later question.
Ideally, the greatest example would be a piston-prop from this era made in two versions: land and naval. Then we could just compare loaded weights. But I can't think of any. Don't think I've ever heard of that. All of the carrier-based aircraft from WW2 that I know of, were purpose built for carriers and never had a land version. Hopefully someone knows where to find airframe specs for these aircraft and has the engineering know-how to say how much steel would be unnecessary for a land version.

Related: https://aviation.stackexchange.com/questions/30061/why-was-the-p-51-mustang-not-adopted-by-the-u-s-navy

The British adopted the Spitfire and Hurricane for carrier use.
Using the Spitfire as an example, here's data from Wikipedia.
Spitfire Mk.VB is 5065lbs clean.
Seafire Mk.III is 6204lbs clean.
They're not exactly equivalent otherwise, but close enough.
So the Seafire is roughly 1140lbs heavier than the land based equivalent, or over 20%.
Of course this will vary wildly for different variants compared, but it's a ballpark figure

@ironduke97 OP asked specifically about WW2 prop aircraft, last I checked the F-4 doesn't match either criteria :)

oh my bad, I wasn't paying attention.

Well well well, the Spitfire was adapted for carriers? Was that ever used in the Pacific Theater against the Japanese? And wow, 22% heavier... That is, huge. I hope someone else can find a corroborating example because this is just, huge.

@DrZ214 Wikipedia says they were used in the Pacific theater from 1944, so yes. and not just in WW2, but the Korean war as well

Also take into account that the extra weight is caused by the double-hinged folding wings and the more powerful engine. It's not only because of airframe strengthening.

@MadMarky Yeah I was wondering about that kinda stuff too. How much does a tailhook weight? How much does the folding wings component weight? And really, are there different (heavier) materials used to resist saltwater corrosion? But I was afraid of being too broad so I was planning to ask those separately.

@jwenting I found a conflicting source for the empty weight, because wikipedia is not being self-consistent again. https://en.wikipedia.org/wiki/Supermarine_Spitfire_variants:_specifications,_performance_and_armament That page shows many variants and their weights. The Seafire Mk.III empty weight on that page is much less than the main wiki page for the Seafire. Interestingly, the Mk.XV is much closer to 6204 lbs. I would guess the main page accidentally took the numbers from the wrong variant. However, both pages have sources so it's impossible to know which one is wrong (or both).

Anyway, if you do the new math, the Seafire Mk.III empty weight is 7.6% more than the Spitfire Mk.VB. Maybe a structural engineer can comment on how more believable, or not, this is to the original figure of 22%.


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What is the Fuel-Oxygen ratio for a large turbofan at cruise conditions?
What is the Fuel-Oxygen ratio (by mass) for a large turbofan at cruise conditions?
In order to be concrete, I'll just pick an engine at random. I pick the GE90-115B. Okay it's not really random, it's just the largest turbofan I know of.

for what its worth: the ge90-115b is the largest turbofan (it has the world record for most thrust)

I chose the GE90-115B because it is the largest turbofan in terms of thrust. Unfortunately, it was hard to come by specific fuel consumption figures for cruise conditions, maybe because the -115B is a relatively new version of the GE90.
But I finally found enough for a preliminary answer, so here it is.
From this pdf, I found that the GE90 at cruise consumes 1.079 kg/s of fuel. It also cites an air-flow of 576 kg/s. But I am not sure how trustworthy this pdf is, as it does not specify the exact model of the GE90 (there are many) and it had no in-line citation, and the only web link in its list of references was "http://www.ge.com/geae/ge90" which is a dead link. Nevertheless it is the closest thing I've found so far for a GE90-115B's fuel consumption at cruise.
(I did try to investigate GE's official site but it is hard to find specific data, which is a real shame.)
From Wikipedia's specs, I find that the 777 cruise speed is 892 kmh and service ceiling (which I'm assuming is cruise altitude) is 13.1 km.
From more Wikipedia specs, I find the fan diameter is 330 cm.

Therefore I can calculate the air intake via the cross section ($\pi \cdot 1.65^2$ = 8.553 sq meters) and the airspeed (892 kmh = 247.778 m/s) and the air density at 13.1 km above MSL from this chart (which I hate btw because it should put altitude on the x-axis, but instead it's on the y-axis, making it very hard to read), which I read as 0.2325 kg/m$^3$.
So air flow rate is 247.778 * 8.553 * 0.2325 = 492.725 kg/s.
The pdf says air flow is 576 kg/s. I'm not sure if this is a significant disagreement. In this case, I think it is reasonable to believe the pdf because the air intake cross section should be a little bit bigger than the fan diameter. The nacelle cowling is curved, so the diameter we want should be from the axis to the midpoint of the cowling. Judging from this photo, the extra radius looks like about 13.5% more, so 1.1335^2 = 1.2896 and multiplying that by 493 gives us 635.773 kg/s, only slightly closer to the pdf's value.
Anyway if we use 576 kg/s of airflow and 1.079 kg/s of fuel flow, we would have a fuel-air ratio of 1:534.
But that can't be right, seems way too high, and indeed there's more. We need to multiply by the bypass ratio...actually the inverse of it (and off by 1) because we only want airflow through the core. Then multiply by 0.2 because oxygen is only about 20% of air by mass.
The bypass ratio according to this is 9:1. But according to this pdf it's 7.08:1. Personally I think 9:1 is more believable because the GE90-115B is supposed to be a recent, advanced turbofan with the latest technology, but at the end of the day that's only my guess. If anyone can reconcile these disagreeing figures and show which one is wrong, please let me know.
Edit: Bypass ratio in cruise, according to the original pdf, is 8.1:1. Thanks again to Peter Kampf for clarifying that this figure looks realistic. I adjusted the math accordingly.
Therefore, 534*0.2/9.1... I get a fuel-oxygen ratio of 1:11.7 ! For simplification we could call this a 1:12 ratio of fuel to oxygen.
Interesting. This is not stoichiometric. Wikipedia says the stoichiometric for dodecane (C12H26) and O2 is 1 mole to 18.5 moles, so if we assume the mass of dodecane is 170 and the mass of O2 is 32, then the stoichiometric ratio by mass is 170:592 = 1:3.482 !
(Btw, dodecane is the simplified chemical for kerosene, which itself is the typically-assumed formula for jet fuel.)
So the 1:12 ratio is very oxygen-rich. I don't know for sure how to interpret this. I guess jet engines just cannot extract more power for some reason? I remember they are limited by core temperatures to prevent the blades from melting. That could be what's preventing them from burning stoichiometrically.
There is something else missing from the puzzle too. Is fuel combustion complete? That is, are all fuel molecules burned or are some just not burned for some reason as it goes through the engine? I know photos of older jet engines tend to have a lot of smoke, which means a lot of unburnt (wasted) fuel. Today I don't think I've ever seen smoke from a big turbofan, so I'm gonna guess that unburnt fuel is so low that we don't have to worry about it in our F-O ratio.
Edit: thanks to Peter Kampf for confirming, burn-rate is >99% and the less than stoichiometric ratio is indeed because temperatures would otherwise be too great for turbine materials of today.

That was a lot more work than it should have been, and if we're being honest here, it's not done yet because I would really like clarification on the 1.079 kg/s of fuel flow the pdf cited. It did not specify which version of the GE90 engine it was quoting. It's been pointed out that cruise conditions vary so there is no one authoritative figure. Too true. I guess the most authoritative figure would be what the manufacture designed it for at cruise conditions and maximum load right after it reached cruise.
If I had known how much trouble it would be to find this figure, I might not have chosen the GE90-115B. But it is the largest engine by thrust, so I think it's an important figure to know. I hope I did it justice.

Bypass ratio is not constant, but changes with flight speed. Same goes for SFC (it doubles between SLS and cruise). You always need to make sure that the conditions for any figure are well defined. The reason for the oxygen-rich combustion is the maximum turbine temperature: If burning would be stoichiometric, the turbine would melt! Not only nitrogen but also oxygen are used as a process gas. The oxygen-rich, hot combustion is the reason that NOx emissions have gone up, like in modern diesel engines.

Advances in burner design result in combustion efficiency of >99% (that is also quoted in the PDF). So yes, combustion is nearly complete.

@PeterKämpf Ah, so the ratio is 9:1 at takeoff but 7.08:1 at cruise? And what does SLS stand for? Thanks for the stoichiometric and burn-rate clarifications too, I will make an edit.

SLS is sea level static. This is the performance on a test stand or on the ground. The change in the bypass ratio is not as wide as your numbers indicate. The figures in the PDF look more realistic (8.4 at take-off and 8.1 in cruise).

@PeterKämpf Thanks again. I edited and adjusted the math accordingly. New ratio is 1:11.7, which fortunately doesn't really change any conclusions.

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How much electrical power does one generator produce on a large turbofan?
From my related question, now I want to know just how much power does the generator produce? In other words, what is the max rating of the generator? It looks so small and that surprised me, but since it spins very fast it could get high power in a small package.
The generator from any large turbofan is fine. The answer will also help us figure how much electrical power a large commercial airliner could need at any given moment.

This question is really too broad and dependent upon the engine in question.

Keep in mind there is a close correspondence between copper mass and kVA, and a close inverse correspondence between iron (magnetic) rotor mass and frequency.  Hence the high frequencies of aircraft really reduce the iron mass.  Nothing can reduce the copper mass.  Meanwhile railways love 16.67 and 25 Hz because weight translates into traction, so they don't mind the iron mass.

There are several max ratings of a particular generator, depending on the duty cycle:

peak rating (5 sec);
max. Continuous (no time limit);
and often an intermediate value for 5 minutes is given.

The generator produces internal heat together with the electricity, which is not easy to remove by forced cooling and must dissipate through convection. The generator is rated in kVA, not kW, because the three phases consist of sine waves which are not in phase with each other. The kVA rating denotes the electrical power used, a kW rating is for the useful power output.
The Integrated Drive Units consist of a drive mechanism and the generator itself. The generator is a fast spinning device, rotating at a high, constant velocity. Since the rotational speed of the LP shaft is not constant, a constant speed mechanism is added between the generator and the LP axis. Classic technology is a hydraulic-mechanical solution with associated losses (like the automatic transmission in a motor car), newer technology such as presented here uses Constant Speed Variable Transmission.

The generator example in the picture is typical for single aisle and many twin aisle aircraft, and has the following power rating:

Cont.Rating : 90kVA
5 min. Rating: 112.5kVA
5 sec. Rating : 150kVA
Frequency:400±5Hz(3)
Voltage:115/200V
Size:375L×460W×330Hmm
Weight:54.0kgf(dry)

Three IGTs of this class make up the AC power capability of a typical single isle aircraft such as the A320, the third generator being driven off of the APU and normally not operating during cruise. The A380 uses six generators of this same rating.
Power demand for some systems scale up with aircraft size. This site gives electrical power demand of a typical classic wide-body jet (pre-bleedless-air):

In-flight entertainment and galley demand scales up with passenger count (33%).
Air conditioning scales up with cabin volume (19%).
Fuel system demand scales up with engine size (10%).
Wing anti-icing scales up with wing span (8%).

Other electricity users such as the avionics are a relatively constant factor as a function of aircraft size. The same site gives the power demand as a function of flight time for a single aisle aircraft, showing that one 90 kVA generator can drive essential loads plus utility loads. The dotted lines in the graph are for continuous rating.

A380
The A380 has six AC generators on board: one driven by an engine each, and two APU generators. From the A380 FCOM:

Each engine has one generator. These engine-driven generators are the main source of electrical power. When an engine is running, its generator provides 115 V AC power at variable frequency. This frequency ranges from 360 Hz to 800 Hz, depending on the N3 rotation speed of the engine.

And

At least two engine generators (or APU generators) are necessary to supply the entire network.

B787
The B787 has a different arrangement. The B787 has no engine bleed air, and systems normally powered by bleed are now powered by AC electric power. A bleedless aircraft requires considerable extra generating capacity in order to power Cabin Air Compressors (about 19% of total power load for a wide-body) and electrical wing anti-icing systems (about 8% of power load), and the Boeing 787 incorporates four engine-driven Variable Frequency Starter-Generators (VFSGs) of 250 kVA apiece in order to accommodate these loads.

Three such IGTs make up the typical AC power capability of a typical single isle aircraft, such as the A320. Three generators for a twin-engine A320? Where does the 3rd generator go?

The third one is powered by the APU.

This is one of the four Variable Frequency Generator (VFG) of the Airbus A380. It's apparent power is 150 kVA.

Rotor, source: Safran
The total power available from the engines is 600 kVA.
Seen from the other side:

Stator, source: Thales
The A380 is a one of these "more electrical" planes, where hydraulics tends to be replaced by electric devices, and the electric circuits have been optimized by electronics and variable frequency. There are secondary sources for electricity: APU (auxiliary power unit) and RAT (ram air turbine). Overall view from the book Civil Avionics Systems (Moir, Seabridge, Jukes):

Source: Civil Avionics Systems
In case of total failure, the ultimate RAT backup is totally electric-based contrary to conventional aircraft where some hydraulics is also provided by the RAT to move the control surfaces.

Interesting. Why do they measure it in apparent power? 150 kVA should be equivalent to 150 kW, since power is just volts times amps.

Power and apparent power are only equivalent when the load is a pure resistor. When it's a general impedance, that is when there is a capacity or inductance component, then voltage and current don't vary in phase, one is delayed (phase difference $\phi$), and instantaneous $u \times i$ is therefore different than when in phase (there is a real and an imaginary part appearing in the product). See Wikipedia. Actually the relationship is $P=U \times I \times cos \space \phi$

@DrZ214 Or harmonics, or other non-linearity. Power factor is rarely equal to one in AC systems.

150kVA is hardly 120-135kW in general real life

@aroth so do electricians.  Unfortunately, lag and harmonics crashed the party, and all of us have to deal with it.   If you draw 150kW from your 150kVA generator, expect magic smoke.

@DrZ214 because the actual limits of pretty much any kind of generating equipment are going to come at some number of VA. If you give it a rating in W, it has to come with a big fat asterisk, but if you rate it in VA, it will deliver rated VA under specified conditions.

@aroth: "Everything should just rate in kW": Actually 1 kVA = 1 kW, it's just this kW (kVA) doesn't measure the same power than the one measured in kW. VA expresses active+reactive power, while W expresses (for general public) only active power. If you size a wire for one W of AC active power, it will burn if you use it to feed a 1 VA AC motor because of the additional reactive power created by the engine and dissipated as heat. Power means something only if we know what we are talking about. Expressing power in VA vs. W helps understand which one we want to talk about.

Naive question: Assuming the "limits" of generating equipment are anyways some kind of "smoke" i.e. Power; why do they occur on the margins of a VA operating envelope instead of a Watts envelope?

@curious_cat: When an empty capacitor starts to be loaded you have a peak of current but the voltage is null. $P=UI=0 \times I = 0$.  However energy is transferred from the generator to the capacitor as the capacitor can later lighten a led. As there is current (with no voltage, but current nevertheless), ohmic loss occurs in the wires in the form of heat (power in heat is $P=RI^2$). So here you can heat a wire with a null active power, only reactive power. Apparent power is the sum of both, and that's what counts here. If you want more details, ping me from the chat.

@mins The way I understand how the terminology is used, is energy input, of which a certain portion is dissipated as heat, and the remainder is useful power output. I've not heard of the term Reactive Power before.

@Koyovis: "I've not heard of the term Reactive Power": Impedance = resistance + reactance (complex sum). Reactance is the sum of capacitive reactance (capacitance) and inductive reactance (inductance). Reactive power is the part associated with the reactance. Reactive power is usually expressed in VAR (VA reactive). Apparition of reactive power denotes the fact that voltage is not in phase with current (said otherwise: power factor < 1). See this article.

all_licenses: Creative Commons - Attribution Share-Alike - https://creativecommons.org/licenses/by-sa/3.0/


text:
Where is the aerodynamic center for an A320?
Aerodynamic center, aka the Center of Lift or Center of Pressure. I would like to know where this is for an A320 airplane.
Sadly I don't know how to estimate aerodynamic things. I can do it for the center of mass, but when it comes to aerodynamics, I cannot simply find the geometric center of multiple arranged bodies, each one assumed to be uniform density. If anyone knows a trick or method of estimating the aerodynamic center, please let me know.

The location of the aerodynamic centre can be tricky to find in an existing aeroplane, it will shift as a function of speed, Angle of Attack, flap deflection etc. In the pre-design phase the aerodynamic centre is often taken at 25% of Mean Aerodynamic Chord, and since you draw a plan view of the aircraft, we can use this to get an approximation.
If the wing has straight taper over the planform, the MAC can be constructed like in the picture below. Then take the point at 25%: this approximates the location of the aerodynamic centre for one wing, the other side is identical of course, and aerodynamic centre of both wings is halfway the connection line of the two.

 However, like many transport jets, the A320 has a trailing edge sweep of zero next to the fuselage to accommodate the main gear. So the aerodynamic centre shifts a bit aft. You can construct it or estimate it, aerodynamicists may frown upon this. There are many sites that show the graphical method of the picture above (here's mine), because it is nice and simple. To illustrate the complexity of finding the MAC in other cases, here is a graph from the same source.

Update
The above method gets us the MAC, which is the first step in determining the AC of the aircraft. A further method for finding AC in the pre-design phase is presented in Appendix E of Torenbeek:

Find wing AC as above.
Wing/Fuselage pitching moment, Appendix E.7 This is sensitive to the pressure distribution and is therefore hard to predict, except for:

Forward shift due to fuselage sections fwd & aft of the wing, particularly the nose. Based on the ratio of fuselage lengths.
Correction for lift loss in the wing/fuselage lift carry-over according to DATCOM.

Fuselage pitching moment. Equation obtained from Munk's theory for near circular cross sections.
Nacelle contributions. Appendix E.8, cannot be predicted reliably other than by wind tunnel tests, only a statistical correction is given.

The above is used for determining the AC of the aircraft. The contribution of the tailplane is included in the lift equations since it is dependent upon the Angle of Attack of the aircraft.
The complete method including graphs and equations is too lengthy to present in this forum, and can be consulted directly in the referenced Appendix.

Thanks but this seems to only be for the wing. The AC of the whole airplane has to take into account the 3 tail surfaces too. Also, I'm not sure I understand the first picture.There's a circle marking the quarter chord point in the middle of the span, but it also constructs a big X that intersects lower. Where is the AC for the wing? Is it the circle or the X?

The big X is for constructing the mid point of the Mean Aerodynamic Chord, the AC of the wing is then taken at 25% of the MAC. Yes the wing/body blended area and the fuselage have influence on the MAC as well. The tail surfaces can be trimmed, horizontal tail deflection is used to balance shifts in CoG. The wing cannot be trimmed, so in the pre-design stage the AC of the plane uses wing & fuselage geometry for determining the dimensions of the aircraft. But of course you're right, the rest of the aeroplane must be considered as well.


text:
Are there turbofan/turbojet analogues to a centrifugal-compressing jet engine?
Axial-compressor jet engines fall into two categories (and only two as far as I'm aware): turbofan and turbojet. I'm wondering if centrifugal-compressor jets also have such a distinction or something similar.
I cannot imagine how a big rotating fan-like object could both compress air centrifugally and bypass it out the rear, but I would not be surprised if some engineer found a clever way of doing it. And anyway, a centrifugal jet could be separated into two broad types based on something slightly different.

I don't understand the question. Are you only interested in the bypassed air and asking if there are engines with a centrifugal fan (whatever the core solution) where the cold flow is accelerated by a centrifugal solution?

@mins It's more about bypassed air, whether or not there's a fan doing it. Would be interesting to see if the centrifugal compressor could do that on its own. Centrifgual compressor jet engine with relatively high bypass ratio, 2:1 or more, if such a thing exists.

Are we ignoring turbo-props? If so, why?

The fan in a turbofan engine is nothing but an air compressor, and does exactly the same as a centrifugal compressor. The axial arrangement in a turbofan allows for neat separation of bypass flow and hot flow, and this is the only (relatively simple) engineering problem that would have to be solved in a centrifugal compressor engine. Then, just blow some of the compressed air out into the atmosphere, via a suitable exhaust shape.
Axial engines allow for straight through flow, without having to bend it around corners with associated losses. They are more compact though, and therefore mostly found in turboshaft applications where jet exhaust power is negligible.


text:
What’s service does ATC service include?
As I’ve studied on ATC service and airspace class by referring to ICAO Annex, I have some doubts on ATC service.
According to ICAO Annex 11, Class D airspace is defined as “IFR and VFR flights are permitted and all flights are provided with air traffic control service, IFR flights are separated from other IFR flights and receive traffic information in respect of VFR flights, VFR flights receive traffic information in respect of all other flights.”
When referring to other definition such as ATS service, It is divided into ATC service, flight information service and alerting service. And I think separation and traffic information service are included to ATC service of ATS service.
So, when I read the definition of Class D airspace that IFR and VFR aircraft are provided with ATC service, I think separation and traffic information service are also provided because these are the ATC service. However, as you guys know, although ATC service is provided to all VFR and IFR in controlled airspace, separation and traffic information are selectively applied depending on airspace class.
To sum up, I want to know what ATC service includes and whether traffic information and separation are one of ATC service or not.

Air traffic control service is a service provided for the purpose of:

preventing collisions:

between aircraft, and
on the manoeuvring area between aircraft and obstructions; and

expediting and maintaining an orderly flow of air traffic.

(ICAO Annex 11)

When saying a flight is provided with ATC service, it is just another way of saying that it is a controlled flight. That is, a flight where the pilot can't just do what they want, because the flight is subject to an ATC clearance; the ATCO will clear the flight to follow a certain route and fly at a certain level.
Importantly, all controlled flights are coordinated as long as they remain in controlled airspace. This ensures alerting service throughout the controlled portion of the flight. This coordination also extends to military units (so we don't have a repeat of Korean Air 007). Controlled flights are also provided with flight information service (traffic info, weather info, info on nav-aids etc.)
As stated in the quote above, ATC service is provided for the purpose of preventing collisions. However, actual separation is not necessarily guaranteed to be provided by the ATCO. The class D airspace you mention is a great example. Separation is only guaranteed to be provided by ATC for IFR-IFR conflicts. In other cases, it is ultimately the responsibility of the pilot* to prevent collisions, which they are able to do based on the flight information service provided.
*: the controller can still decide to provide separation, but is not required to do so.
Note that separation on the manoeuvring area (which includes the runway) is provided regardless of airspace class.

Thanks! I understand clearly!


text:
Is GLS (GBAS landing system) an RNAV approach?
PBN which is contrast to conventional navigation consists of RNAV and RNP. And I know that general RNAV approach of IAP has up to 4 minima such as LPV, LNAV, LNAV/VNAV, and LP.
And I know that GLS approach is precision approach that use GBAS as well as GPS, but GLS approach procedure chart is separated with RNAV approach chart. I guess that GLS approach is RNAV approach because it use GPS system not conventional aids. Is it correct? is there any official document?


text:
Can the fuel on the wing cause the aircraft to destabilize?
If fuel is stored in the wings, which are in constant movement during the flight, can the fuel unbalance the aircraft?

There are baffles inside to minimize sloshing.

Yes it can. Source: I'm a pilot.
Most aircraft (even many small general aviation ones) have balance parameters where the fuel quantity in each side cannot/should not exceed that on the other side by greater than a certain amount.
Note that this is usually caused by fuel consumption in flight, not sloshing as you asked.
Also, (as in your question) fuel can slosh between tanks in most aircraft through the vent line that often connects one tank to another such as in light Cessnas. However this is usually only a small quantity and furthermore the aircraft is usually in level flight so it's not really a problem.
Fuel sloshing inside of one tank could potentially imbalance the plane a bit but I've never ran into any real effects of this nor heard anyone mention it. Furthermore, most aircraft have some sort I baffles in the tanks (usually wig ribs) that prevent much sloshing, whether by design or not.

For example, Airbus A320, for fuel umbalance, 300 kg.


text:
How does an airline open a new route?
How do they manage to open a new route? For example Dubai-Panama. Do they have to open new airways?

If you've never driven from Paris to Rome, do you need to build new roads?  It's exactly the same.

@Simon Your analogy is broken. Nobody has ever flown from Dubai to Panama before. If nobody had ever driven from Paris to Rome before, would I have to open new roads? I've no idea. Maybe, maybe not.

@DavidRicherby It's not broken.  There are airways in and out of Panama.  There are airways in out of Dubai.  You just need to join them up into a route.  If there is a break, then you need a new airway.  There are roads in and out of Somewhereville.  There are roads in and out of Otherplaceville.  You just need to join them up into a route.  If there is a break, then you need a new road.

@Simon "If there is a break, then you need a new road." Well, gee, how do I know if there's a break or not? Is there are a break? Maybe somebody wants to know that. Maybe I should post a Stack Exchange question about it. Oh, wait. Somebody already did. You seem to be criticizing the question based on the fact that you know the answer to it. Well, other people don't know the answer, which is why one of them asked.

@DavidRicherby I didn't criticise anything.  I simply offered an analogy as a comment knowing that someone would proffer a good answer or two.  It's a good analogy too.  In all of the basic meaningful ways, when asking about airways, flying from A-B is the same as driving from A-B.  Are there roads?  Do I have the right maps?  Is there a route etc etc.  I offered an analogy, not an answer, and criticised precisely nothing.

@Simon paris to rome, in just 80 minutes

@Simon I'm glad you intended no criticism. Given that your comment essentially says, "This is identical to a situation that everyone is familiar with" it looked like criticism to me.

The road analogy also breaks down in other ways. If I need to build a new road between cities, generally my only option is to convince the government bodies involved that they should build the roads; I can't just do it with private funds. And also, if I want to drive my car on existing roads, generally speaking I just do it; I don't need to go through any regulatory body. It's not obvious to me that either of those aspects, which are a huge part of driving from one city to another, applies to airlines.

Related: Freedoms of the air

@Simon You can travel by train

@yshavit There should be a badge for pedantry.  If you want to drive, you must have a license, you must have insurance, you must have paperwork (or international agreements in place) to cross borders, you must have a mechanical fitness inspection on the car. If you need a new airway, you must convince government bodies  You can fly where there are no airways, you can drive where there are no roads.  It's just a simple analogy, and I did say in the basic meaningful ways.  It works.  I'm really not looking for a pat on the back, simply offering a way to think about it.

No. You do not need to create new airways. You stick up to the already existent waypoints or not, if the route flies through a segment of free-route airspace.
Also, the engineering department of the airline studies the route, for instance: computes the MSA (Minimum Safe Altitude) for every point, engine failure and depressurization simulations are done in order to avoid possible hazards in critical points, if the aircraft is ETOPS certified they will decide which alternative airports are available for the route, fuel computations, chart generation for arrival and departure (most of the airlines use their own charts instead of those provided by the ANSP), et cetera.
Each airline has its own policy and the process may vary but fundamentally the procedure done is to make a study of all the parameters that can affect to guarantee safety and efficiency.

In addition to Airman01's answer, opening a route goes beyond flying from point A to point B, you also need to:

get landing and departure slots at each of the airports. At some with lots of spare capacity, this is easy. At others, such as the largest and busiest airports, this is very difficult, and is a process that can sometimes take years. Airlines have been known to buy competitors just to get their slots.

arrange ground handling (jetways or stairs, loading/unloading luggage, refuelling, possibly catering and light maintenance, check-in and gate personnel). Sometimes this may be provided by the airport itself, sometimes it will be outsourced to specialised companies doing so in many places, sometimes to local companies, and other times it will be handled by the airline itself.

have all the relevant licenses/authorizations from countries overflown as well as source and destination, and arrange for the payment of relevant taxes and fees to the appropriate authorities.

and then you need to insert this into the rotation of the planes and crews in order to minimise downtime while still trying to find the most convenient times for your customers.

EDIT
I forgot to mention that in the case of international flights, there's the issue of the Freedoms of the Air. Overflying, stopping, flying to, from or between countries other than your own is not automatic, and is usually subject to state-level agreements.
Overflying another country (First Freedom) is nearly universal (though the two states must either both be signatories of the IASTA, or have a bilateral agreement). The Wikipedia page on the topic also lists examples of flyover fees.
Further freedoms are a much more complex issue, though relatively recent changes like the EU–US Open Skies Agreement changed things a lot on many of the most competitive (and restricted) routes.

And ATC in the related and adjacent FIRs?

Maybe this is an exception which proves the rule, or maybe this is not exactly a valid answer because it doesn't relate to the design a new complete route, but sometimes new airway segments are created, e.g. this discussion between IATA and FAA for opening two new polar segments:
Summary of Discussions of the Sixteenth Meeting of the Cross Polar Trans East Air Traffic Management Providers Working Group (CPWG/16) 3-6 December 2013 – Ottawa, Canada

8.1. Based on ideas and proposals from IATA member airlines, State ATM Corporation developed two new route proposals that would supplement
  the existing cross polar routes and provide additional flexibility.
  The routes presented were:
a. 7957N16858W – RODOK  to join G495
  b. 7457N16858W – LUTEM – OLMIN – TURAN – ASKIB
8.2. The FAA agreed to review the routes and make an effort to limit restrictions.  These routes will be added to the ATS Route Catalogue.

The routes were later discussed at ICAO, but seems to be still waiting for their implementation. However in the same document other routes are adjusted in the Sea of Japan:

I'm surprised that not only is there a Pyongyang FIR, but it has routes going through it.

@HenningMakholm: Agreed. You use it at your own risks.

Cross Polar Trans East Air Traffic Management Providers Working Group that's nice and concise...


text:
If the engine of a helicopter fails, is it possible to land safely?
Imagine the rotor of a helicopter is not spinning anymore, can you control the aircraft to land safely?

To answer your question directly, if the rotor is not spinning, then no.  You will crash.  I suspect that's not what you meant though.  If the engine stops, this does not mean that the rotor stops.  As long as the rotor is spinning, in the correct range of RPM, then yes, you can fly safely.

Like the propellor on an airplane, the rotor is actually not there for lift or propulsion, it is for keeping the pilot cool. You should see him sweat when that thing stops turning...!

You're asking two different questions, actually.

If the engines fails in a helicopter, the rotor will usually rotate and the pilot can land the helicopter by autorotation. The record is from an altitude of more than 12km by Jean Boulet when he descended via autorotation as the engine flamed out and couldn't be restarted.
If the rotor seizes for some reason in a helicopter, there is no lift produced and it will fall down. For tilt rotors like V-22, if the conditions permit (if it is in aircraft mode), the aircraft can be glided like normal fixed wing aircraft.

Helicopter can land via autorotation.

This is a bad diagram.  You do not dive to maintain speed and rotor RPM.

@Simon so what do you do to maintain speed on the rotor?

@kepler22b Initially, you lower the collective, to remove the pitch and therefore the drag, then flare to cause the blades to cone and increase the speed of the disc.  After that, you fly normally, albeit with a high rate of descent, until the flare at the bottom.  During the main part of the auto, you control speed with attitude, just as in normal flight, and control rate of descent with speed. If you dive, the rotor speed decays.  If you dive at the point where the engine quits, you can get into big trouble quickly.  Lever down, nose up.

to maintain the speed of the rotor, you maintain slightly aft cyclic so that he forward motion of the copter goes up through the driving part of the blade.


text:
Why do the spoilers move in this video?
In this video when the aircraft is turning left(2:25) the spoilers on the wing move. What causes them to move?

See this: http://aviation.stackexchange.com/q/140/520

Two reasons that are really one reason. In short: Because lift.
In this case though, it's "to get rid of lift". The spoiler's job is to reduce the amount of lift the wing experiences. This can be done symmetrically, on both wings at the same time, to increase the rate of descent or keep an aircraft from lifting back off the ground when in ground effect, or in this case, it can be done asymmetrically, or only on one wing.
In order to turn, an aircraft has to create an imbalance of force, more on the side outside the turn, less on the inside of the turn. The wing to the inside of the turn needs to drop, but it doesn't really want to...it has lots of air pushing it up (when it doesn't, you will have a very bad day)!
The spoiler in this case is just acting as an extension of the aileron. By moving upward with the aileron itself, it spoils the airflow over the top of the wing a bit, reducing the lift, and causing the wing to drop a bit, raising the other wing. It also acts as a larger surface area for the incident airflow to impart force on, which also pushes the wing down.
On the other side, the aileron is doing the opposite, increasing the wing's effective camber, at least near the tip, and increasing lift. Flaps do the same, but also are usually used symmetrically to increase lift on both wings. In some airplanes, you can have a 'flaperon', which can also be used as an additional control surface while also doing it's flap-thing.

this sort of answers the question, but not really. Most aircraft use opposite aileron on the inside wing, but as you can see in the picture, aileron is in its normal position. The question in this case is: why is this plane using the spoiler to reduce lift on the inside wing instead of the aileron? the answer is in this question: http://aviation.stackexchange.com/questions/140/is-roll-control-achieved-exclusively-with-spoilers-on-an-a320-on-final-approach

That is controlled by the pilot (indirectly via the computer system) who uses the spoiler to control the turn of the airplane instead of ailerons. Notice that the airplane turns when the spoilers move. It does so by killing the lift on one side by "spoiling" the airflow while the lift on the other wing creates the roll and bank moment with the "unspoiled" lift.

no, its done automatically

The pilot does not directly directly control asymmetrical spoiler deployment. In the case of the aircraft in the question, the flight control computer commands asymmetrical spoiler deployment based on it's interpretations of the autopilot or pilot control commands combined with other flight data inputs.

that is a very literal answer to the question!


text:
Can a helicopter stall?
Imagine that a helicopter raises its nose up and then there is a loss in speed. Can it enter a stall? If so, how do you recover from the stall?

By the way, raising the nose up and loss of speed is a typical entry to a hover from forward flight. A helicopters forward speed does not contribute to a stall like it does for fixed wing aircraft unless you are just talking about the retreating blade stall. At zero forward speed the retreating blade hits the relative wind just like the advancing blade, so it won't stall. Moving forward the retreating blade is traveling slower (relative to the air) so it can stall.

The simplest answer is yes, but not how you might think.  There are no useful comparisons between stalls in fixed wings and stalls in helicopters.

The blades can stall if the AOA is high enough, this is known as a rotor stall. See the link that @Simon provided for more information Wikipedia article.

@Simon what is OP?

@kepler22b Sorry, "original poster" or "opening post".  It means you ;)

Yes and No.
The definition of a stall is that the airfoil stops generating lift, which happens when the critical angle of attack is reached.
In an airplane, this happens during normal flight by pitching up until the critical angle of the airfoil is reached.
In a helicopter, aft cyclic ("pitching up" in an airplane) will only serve to make the helicopter climb and slow down until it reaches equilibrium in the new attitude.
Helicopters, however, can suffer from what is called "retreating blade stall," when the blade is at full pitch on the retreating side of the rotor:

This doesn't happen under normal circumstances, but there are four factors which contribute to retreating blade stall:

fast -- close to Vne
heavy -- at max gross weight
hot -- above 35C
high -- at high altitude

Retreating blade stall is considered an emergency, and can be recovered from by reducing power.

Above 35F or 35C?

Is it possible to stall all blades by yanking up on the collective?

no because so long as the rotor is turning at the right RPM, the relative wind is below the critical angle. if the rotor isn't turning fast enough, then you have much bigger problems.

@Steve It might be possible on some models, but it would be the helicopter equivalent to Pugachev's Cobra: stall by extreme pitch up while far away from stall speed.

There is the premise the rotor is spinning sufficiently fast. There is the other kind of stall when the rotor RPM drops too low.

A stall, as in slowing down and then descending too fast - yes, in the following scenario:

Helicopter flies with forward speed, above the hover ceiling.
Pilot moves cyclic to aft, and pulls up collective to maintain height.
When all airspeed is gone, the helicopter cannot maintain hover at this weight/altitude combination and starts to lose altitude.
Descending vertically, the helicopter descends in its own downwash and enters a dangerous situation called Vortex Ring State. Even if more collective travel was available, it would not create more lift in this state.

There is a discussion possible about what stall exactly is. If it is flow separation, the rotor blades are technically not stalled as @Jan Hudec points out. If it is stopping the generation of aerodynamic lift, this is exactly what happens during both fixed wing stall and rotor wing vortex ring state.
Even the way to get out of vortex ring state is the same as with fixed wing stall: move the cyclic forward so the nose tips down. Or sideways alternatively left/right and descend like a leaf from a tree.

This is technically not a stall, as the flow is indeed still attached. In fact, the air is moving downwards, reducing the angle of attack and that is why the rotor can't provide enough lift (and raising collective just accelerates the circulation instead of providing more lift). It is, however, just as dangerous, if not more, than a stall in a fixed-wing aircraft.

@JanHudec Have amended the answer, thx.

The rotor is moving the air down, at about the same velocity as the helicopter is coming down.

@qqjkztd, when creating lift, the air must accelerate downward (by principle of action and reaction). So below the helicopter there is air moving downward (called downwash). And vortex ring state (described in this answer) occurs when the helicopter descends into its own downwash.

It is possible to slow down, indeed come to a stop, and descend "too fast" (whatever that means) without stalling.  There is only one type of stall in a helicopter which is a blade stall, either main rotor or tail rotor. I don't understand you last sentence - "descend like a leaf from a tree."  Are you talking about auto-rotation?  This is one way to escape VRS if height permits but the standard recovery is to fly out of the vortex in any direction. Moving the cyclic forwards is just one possible recovery. VRS is not a stall condition and in general, your answer does not address stalls.

Are the steps you list supposed to be sequential?   It does not describe how a blade stall occurs. It may be possible to enter VRS with these steps but I would expect rotor droop to occur first requiring an immediate reduction in power required (or a stall will follow) and since you are above the hover ceiling, as you slowed through translational lift, you would start to descend. You seem to be implying that VRS is a stalled condition.  It is not.

The blade profile is not stalled, as stated in the answer. You're correct in stating that, if we define stall as loss of aerodynamic lift by exceeding critical angle of attack, the helicopter is not stalled. However, the rotor disk has loss of lift in VRS and the symptoms are very similar to fixed wing stall. I am assuming that that is the background of the question.

But the question is quite simply "can a helicopter stall". Your answer does not address this.

How about the first sentence of the answer? I believe I've made a useful comparison with stall in a fixed wing.

Forward motion is key to auto-rotation.  Autorotation with proper blade pitch maintains lift.
In that sense powerless helicopters can be considered "gliders".  Both pitch down to utilize gravity to maintain forward airspeed.  Helicopters (and autogyros) indirectly maintain rotor speed by converting differences in drag from forward airspeed to angular rotor velocity, just like an anemometer.
If a helicopter loses power, it will begin to sink.  Similar to a glider, pitching down is key maintaining controlled flight.  The Bell UH-1 Iroquois has a horizontal stablilizer structure on its tail to assist this manuever.

Can a helicopter stall?

Lack of understanding and proper training can certainly lead to a very bad situation if power is lost.  If one pulled up into a hover$^2$ with the blade at full angle of attack, lost rotation, and started sinking, the entire rotor could stall and may even start rotating backwards$^1$.
But even a helicopter without a horizontal stabilizer pitches down in a sink if the tail produces more drag torque around the center of gravity than the nose.
$^1$ unless blade pitch is brought to negative
$^2$ with sufficient rotor angular momentum, this manuever could be done for a limited amount of time - such as just prior to landing.

Hello I wanted to give this question an answer that is technically correct. The truth is no you cannot stall a helicopter. But yes you can recover from conditions that would normally stall a helicopter before it stalls. Not to plagiarize wikipedia. But lets start with the lift equation:

L=1/2 ρ×V$^2$×Sref×CLift
where:

L denotes lift force.
V defines the velocity of aircraft expressed in m/s.
ρ is air density, affected by altitude.
Sref is the reference area or the wing area of an aircraft measured in square metres.
CL is the coefficient of lift, depending on the angle of attack and the type of aerofoil.
Or you could say Charlie Licked twice on a Pizza under the table and Vomited on Sunday.

Clift = 2 L / (p×V$^2$×Sref)

Lets see how lift is defined
"A fluid flowing around an object exerts a force on it. Lift is the component of this force that is perpendicular to the oncoming flow direction.[1] It contrasts with the drag force, which is the component of the force parallel to the flow direction. Lift conventionally acts in an upward direction in order to counter the force of gravity, but it can act in any direction at right angles to the flow."
Now lets see how an airfoil stalling is defined
"The critical angle of attack is the angle of attack which produces the maximum lift coefficient. This is also called the "stall angle of attack". Below the critical angle of attack, as the angle of attack decreases, the lift coefficient decreases. Conversely, above the critical angle of attack, as the angle of attack increases, the air begins to flow less smoothly over the upper surface of the airfoil and begins to separate from the upper surface. On most airfoil shapes, as the angle of attack increases, the upper surface separation point of the flow moves from the trailing edge towards the leading edge. At the critical angle of attack, upper surface flow is more separated and the airfoil or wing is producing its maximum lift coefficient. As the angle of attack increases further, the upper surface flow becomes more fully separated and the lift coefficient reduces further.[7]"
Or simplified:
"In fluid dynamics, a stall is a reduction in the lift coefficient generated by a foil as angle of attack increases.[1] This occurs when the critical angle of attack of the foil is exceeded. The critical angle of attack is typically about 15°, but it may vary significantly depending on the fluid, foil, and Reynolds number."
So a stall is when an airfoil such as an airplane wing or helicopter blade which is measured from the chord line to the tip is no longer able to generate an airflow where there is less pressure above the airfoil and more pressure below the airfoil generating lift. Some people think of this as the wing has a buoyancy less then the air around it. Not an actually correct definition as to compare to a boat as a boat does not need to move in the water to displace the water less then itself and float above it. But for whatever practical purpose an aircraft moving in the air has less displacement then the air underneath it as long as it keeps moving.
When a wing or wing tip stalls it is no longer creating an airflow generating lift the other forces such as drag and weight are now greater then lift and thrust. In a plane if a plane stalls the plane can regain lift by regaining an airflow around the wings by creating a sufficient angle of attack. Basically when the nose is pointed downwards it regains the ability to generate lift around the wings as the aircraft has regained the ability to create a sufficient angle of attack.
For the angle of attack to generate lift around the blades of a helicopter the blades must be moving and nothing must disturb the airflow around the blade rotor system.
In an auto rotation the engine has failed and the helicopter no longer is producing RPMs from the engine. But the helicopter blades are still moving from the airflow underneath the helicopter. The blades are still producing an airflow allowing the helicopter to slowly float to the ground as though it is doing a vertical descent. As long as the blades continue to provide a correct pitched angle of attack collectively then the helicopter will maintain enough lift to descend the helicopter to the ground safely. Lets see what wikipedia says about an auto rotation:
"Autorotation is a state of flight in which the main rotor system of a helicopter or other rotary-wing aircraft turns by the action of air moving up through the rotor, as with an autogyro, rather than engine power driving the rotor.[1][2][3] The term autorotation dates to a period of early helicopter development between 1915 and 1920, and refers to the rotors turning without the engine.[4] It is analogous to the gliding flight of a fixed-wing aircraft. Autorotation has also evolved to be used by certain trees as a means of disseminating their seeds further.
Airflow through a helicopter rotor. Above, the rotor is powered and pushing air downward, generating lift and thrust. Below, the helicopter rotor has lost power, and the craft is making an emergency landing,
The most common use of autorotation in helicopters is to safely land the aircraft in the event of an engine failure or tail-rotor failure. It is a common emergency procedure taught to helicopter pilots as part of their training.
In normal powered helicopter flight, air is drawn into the main rotor system from above and exhausted downward, but during autorotation, air moves up into the rotor system from below as the helicopter descends. Autorotation is permitted mechanically because of both a freewheeling unit, which allows the main rotor to continue turning even if the engine is not running, as well as aerodynamic forces of relative wind maintaining rotor speed. It is the means by which a helicopter can land safely in the event of complete engine failure. Consequently, all single-engine helicopters must demonstrate this capability to obtain a type certificate.[5]"
Now if the blades were to stop moving the helicopter would stall and drop like a brick to the ground. Once the blades stall completely the helicopter cannot recover.
Another way in which a helicopter can stall is if it descends in its own vortices and these vortices disrupt the ability of the helicopter rotor system blades to create an airflow around the blades generating lift. Lets see what wikipedia has to say about vortex ring state
"The vortex ring state (VRS) is a dangerous aerodynamic condition that may arise in helicopter flight, when a vortex ring system engulfs the rotor, causing severe loss of lift. The vortex ring state is sometimes referred to as settling with power. The Federal Aviation Administration (FAA) sees these terms as synonymous, whereas Transport Canada sees them as two different phenomena.[2]
Vortex ring state, in which airflow is upward on the inner blade section, producing a secondary vortex in addition to the normal wingtip vortices. Turbulent airflow results in loss of rotor efficiency. If allowed to continue, uncommanded pitch and roll oscillations may occur, with a large descent rate.[1]
A vortex ring state sets in when the airflow around a helicopter's main rotor assumes a rotationally symmetrical form over the tips of the blades, supported by a laminar flow over the blade tips, and a countering upflow of air outside and away from the rotor. In this condition, the rotor falls into a new topological state of the surrounding flow field, induced by its own downwash, and suddenly loses lift. Since vortex rings are surprisingly stable fluid dynamical phenomena (a form of topological soliton), the best way to recover from them is to laterally steer clear of them, in order to re-establish lift, and to break them up using maximum engine power, in order to establish turbulence."
The fact is this. A well trained helicopter pilot can recognize situations such as engine failure and vortex ring state before a stall occurs and perform an auto rotation to recover from engine failure or a vuichard maneuver to recover from vortex ring state or to redu collective pitch (to reduce downwash), lower the nose to fly forward out of the downwash, and then apply recovery power to recover from vortex ring state.
But lets be clear. A helicopter can regain from engine failure with an autorotation landing provided the right landscape below and recover from vortex ring state with vortices increasing up the wing tips into the rotor system which will disrupt the airflow around the blades until eventually a total blade stall occurs. A plane can recover from an engine failure by gliding to a landing spot and staying trim. So both a plane and helicopter can recover from a condition which would lead to total stall.
But can a helicopter recover from a total stall the way an airplane can and regain a sufficient angle of attack to regain airflow around the wings? No. A helicopter cannot recover from a critical angle of attack where the entire airflow system around the blades is unable to produce lift. A helicopter is a lot harder to stall then an airplane. It is difficult for an experienced helicopter pilot to mess up a failed engine autorotation recovery or recovery from vortex ring state if he/she is experienced enough then you can recover from these conditions easily by recognizing the situation quickly as the helicopter rotor system RPMs do not decay past the point of no recovery in just a second or so and maintaining RPMs is still possible in an emergency situation.
But if a critical angle of attack is sustained in the entirety of the helicopter rotor blade system and lift is entirely disrupted across the entire blade system and the helicopter drops like a brick like a plane totally nose down in a stalled condition diving to the ground can a helicopter recover? No. The answer is a helicopter is harder to stall then a plane and for an experienced pilot is easier to recover from an emergency with correct maneuvering. But can a helicopter recover from a total stall?
Absolutely not. The lift equation is a theoretical viewpoint based on the principals of flight where of the 4 forces of thrust, lift, drag and weight drag and weight do not overpower lift and thrust and as the 4 forces are equal lift can be sustained. It is very practical. But much like the formula E=Mc2 the lift equation is not a formula that can solve a problem by plugging in numbers for variables. It is not a number cruncher. You have lift or you don't.
Can an airplane recover from total wing stall? Yes it can recover as the correct angle of attack can be reapplied to generate lift around the wings. Can a helicopter recover from total blade stall? Absolutely not.


text:
What is this hole on the tail of an MD-88?
There's a little hole on the leading edge of the vertical stabilizer of this plane:

What is that? What is its purpose?
Thanks
Source: Airliners.net

That is the ram air intake for the air conditioning packs.

At first I thought it might be the APU intake, but other sites say that the APU intake is near the bottom rear of the fuselage.


text:
What is the clearance between a fan blade and the nacelle?
What is the distance between the tip of a fan blade and the inside of the nacelle for a turbofan like the Rolls Royce Trent 900? Does this distance change significantly during cruise due to thermal expansion?

related: https://aviation.stackexchange.com/q/37975/1467

Can't say for sure without specific information from the manufacturer, which would be proprietary, but it's probably somewhere between 0.010" and 0.100". As you make the clearance tighter, you run the risk of a heavy rub between the blades and the case.  But as you make the clearance more open, you lose efficiency.  Somewhere in between there is a sweet spot where efficiency is fairly good and risk of a rub is low enough.  The exact clearance may vary significantly between different manufacturers and models, depending on what type of blades they have.
And yes, this distances changes significantly as you go from idle to takeoff to cruise.  Thermal expansion is a player, but most of it is mechanical loads (i.e. centrifugal force).
also, very minor point, but the clearance you are looking for is between the fan blades and the fan case.  The nacelle is a separate structure that is outside of the case.


text:
How did the gyro gunsights of WW2 get the range and lead of a target?
I just read this article on the gyro gunsight, but I don't understand how it computed range/position to the target and then calculated how much lead is needed.
Surely it must have had radar to get range to the target? But I could only find one mention of radar in the article, and it only pertained to one particular model of the gyro gunsight. Even if so, I don't understand how a radar return is integrated with a gyro instrument.
I understand that a gyro can tell you your roll angle and rate of turn in any or all axes, but the main thing I'm missing is how did it know range and position of the target? It seems only radar can do such a thing and I don't know how radar data could be integrated with a gyro way back in the 1940's. The wiki article even says that these gyro gunsights were developed "just before the Second World War". But I also noticed the article is not that well-cited.

I think the sight has an circular indicator that the pilot can adjust it's size. The pilot adjusts it to match the wing span of the target then the sight would set the distance accordingly.

@mins No I hadn't, but thanks to the answers, now I understand that when wikipedia says "the pilot had to set the wingspan of the target", this means tell it 10 m (or whatever) and keep the reticule at the right diameter. In other words, it needs the absolute size and relative size. Having both of those is enough for a trig calc, no radar necessary.

You are seriously over estimating the technical capabilities of the time.
The gun-sight is not much more than an intricate slide-rule. You had to feed it information about the target size/aircraft type and it used the gyro information from your own aircrafts turn rates to project a reticule circle on the glass.

The position of the projection is offset from the guns aiming point according to the gyro factors, the set range, and the fall rate of the ammo. When the reticule circle lines up with the wingtips of the target you are at the right distance and angle to the target and you push the trigger.
It assumes you are following the same flight pattern as the aircraft in front of you. That is.. chasing it in a dog fight. That makes it fairly limited in use and effectiveness and would be no replacement for a seasoned pilot.
Additional Reference Links:
Aviators Database
Axis History Forum
Video: Gyro Gunsight Mk IID
CORRECTION:
Upon watching the video it is apparent that the implication that the aircraft must be following the same flight pattern as the target is misleading. In reality the target reticule, and in effect the pilot's aircraft, needs to be turning at a rate to track the target's flight pattern. Further, the range can be adjusted while continuing to track the target. That all sounds a bit like rubbing your head while patting your tummy and I am sure took quite a bit of practice and an accommodating target that maintained a constant heading or turn rate in front of you.

When I read "input the target wingspan", I just thought it was something like, push buttons to tell the device that the wingspan is 10 meters. That alone cannot tell you anything about range. Now I see that, in addition to that, the apparent size of the target also has to be input by twisting a dial to make the reticule encircle the target. Both of those things, absolute size and apparent size, are enough info for a trigonometric calculation of range.

@DrZ214 indeed, it is a type of visual range finder. The video does not mention it but if you were catching a target the procedure would be to set the range short so you don't have to keep adjusting it and then wait till the target filled the circle.

@DrZ214 if you notice in the image, setting the wingspan was really just a matter of selecting the target aircraft type on the dial. Popular targets had marks for them. like Junkers-88 (Stuka) , Messerschmitt 110,119 etc.

Other answers here describe the functionality of such gunsights; the "calculation" was done by mechanical analog computer, made up of gears, cams, shafts, and linkages.
Here's part of a Sperry K-3 gunsight computer from a B-17 (inherited from my father, who liked collecting strange gadgets, now decorating my kitchen):

Additional information on the mechanical gunsights and bomb sights of the WWII era can be found here.

Gotta love the true computers :-)

@yo' The concept isn't at all new, either; the Greeks had things like the Antikythera mechanism

I don't understand how it computed range/position to the target and then calculated how much lead is needed.

Reading the Wikipedia article you linked, it didn't, but both assertions are unsourced.
The distance was estimated visually by the gunner through a visual aid:

a reticle to match to a target plane's known wingspan (to adjust the sight for the target's distance)

i.e.: the gunner would know the target wingspan, and its size in the reticle at a given distance, the mismatch would allow the estimation of the distance.
The lead was based on gyro information (hence the instrument name):

It is important to note that the information presented to the pilot was of his own aircraft, that is the deflection/lead calculated was based on his own bank-level, rate of turn, airspeed etc. The assumption was that the flightpath was following the flightpath of the target aircraft, as in a dogfight, therefore the input data was close enough.

Personally, rather than "calculated" I expect that some sort of mechanical linkage was moving the floating reticle around based on the relative position of gyro and airplane.

The video found by Trevor confirms that this is the case.
Starting at 2:00 we are told that:

given the range, the gyro corrects only for the rate of turn
the pilot has to manually adjust the gyro correction for the range (at 4:15 we are told that the renge adjustment does not alter only the position of the reticule, bu also its diameter, to reduce the errors deriving from manually estimating the range)
the pilot can also select the base size of the reticule to match the aircraft being chased (4:35)

The full procedure with the correct sequence of settings and adjustments is described starting at 5:42


text:
Were WW2 American fighters designed to operate from dirt airstrips?
Fighter or fighter-bomber planes built in America during WW2, such as the P-36, P-39, P-47, P-51. Were these designed to takeoff and land on dirt airstrips?
My impression, and correct me if I'm wrong, is the following: Most airbases inside USA/Canada at this time had paved runways. At least half the fighters were designed to takeoff and land on carriers. Also, I think every single photo in wikipedia I've seen of these planes on the ground, is on a paved runway.
Now there is of course the Western Front in Europe, and some American aircraft did get sent to the USSR in lend-lease. I do not know if Britain and France had the luxury of paved runways at the time, but in Russia, most certainly did not. So you would think at least some of them had to had the strong undercarriage required for dirt airstrips. But I am not sure if they came that way or had to be modified after arriving in Europe.
One way to investigate this is to start looking at the airfields America had during WW2. Wikipedia has a giant category page for this, but even tho most of those pages have a history section, sadly I could find nothing about paved/unpaved. However, this photo from Midway shows paved runways, and I'm pretty sure Wheeler Airfield during Pearl Harbor had a paved runway too. If they were doing that in the middle of the Pacific, my guess is that most airbases in the 48 states and east/west coast mainland also had paved runways.
The reason I ask is because a stronger undercarriage requires more weight. Dirt landings require bigger wheels which take up room in the fuselage that could otherwise be used for fuel or bombs. Therefore, if paved surfaces are available, it makes sense to take advantage of it in your design. But as I said above, I can't be sure if paved surfaces would be available everywhere they intended to use them.

American P-51's departing a dirt strip on Iwo Jima The comments about American airbases don't really have any bearing, I don't believe any missions departed airbases from the 48 states other than transport/relocation.

Why do you think a dirt strip needs a stronger landing gear?  Assuming it's graded, of course, and not just a fairly flat field.  From my own (admittedly subjective) experience, a well-graded dirt strip seems a bit more forgiving than pavement.

@jamesqf A previous question of mine, https://aviation.stackexchange.com/questions/19411/how-are-modern-jets-modified-to-takeoff-land-on-a-dirt-runway , asked something like that. The answers I got seemed to indicate that bigger wheels and stronger undercarriage are required.

@DrZ214: But most of the answers there seem to relate to relatively modern jet fighters, and heavy cargo aircraft, not WWII planes.  The relevant factor for propellor planes seems to be wheel loading.  But I think most single-engine planes today (and probably light twins) handle dirt just fine (certainly the Cherokee does :-)), so I don't see why a WWII fighter wouldn't.

@jamesqf, Piper Cherokee weighs 975 kg and lands around 61 knots. That's tiny compared to WWII fighter—P-51 weighs 5,490 kg and lands around 113 knots. And that's not mentioning bombers—even B-17 weighs 29,700 kg (MTOW). While that's still small compared to modern fighters (F-22 weighs up to 38,000 kg!), it's much more than your typical GA SEP or light twin (Piper Seneca has MTOW 2,155 kg, still less than half of P-51).

@Jan Hudec: But it's the wheel loading, not the absolute weight, that matters.  See the link above for pictures of the C-17 and C-130 landing on dirt, both much heavier than a P-51.

@jamesqf, yes, it is. But it was you who made an argument that single-engine planes today handle dirt just fine, which is irrelevant exactly because it is the wheel loading and not weight that matters (and the speeds, which also matter, are much lower).

Early in the war, the RAF was using grass strips. Later in the war, the Allies tended to have prepared runways because they had much better supply and field engineering support than the Axis powers. In the UK, there were special runway teams, that would re-concrete a runway in sections, using fast curing concrete.
For more remote locations, the allies also had Marston Mat, a brilliant invention, the perforated steel planks used to cover dirt runways. Compact for transport, and assembly was very quick with a minimum of tools and equipment needed.
However, in very remote locations where supply wasn't so plentiful, dirt, crushed coral, or grass runways were used. This includes recently captured island bases, and China, which had a difficult supply situation when the Japanese closed the Ledo road, and supplies had to be flown in. My father told me stories about landing on hand built Chinese dirt runways in a P61 Black Widow, which touched down at around 110 kts. A bit bumpy.
Aside from being smoother, easing the load on the landing gear, prepared runways keep the dust and dirt down, which increases engine life. No air filters on those aircraft engines, so the less dust and debris that got sucked in, the longer the engine stayed in service.

In the UK, there were special runway teams, that would re-concrete a runway in sections, using fast curing concrete. Can you link to more details? I would like to read more about that concrete and the procedure used.

The only reference I have comes from a couple of memoir books by bomber aircrew. The US Army had dedicated engineer teams that did nothing but build and repair concrete runways, populated largely by black soldiers (this still being a segregated army at the time). The bomber crews noted that these engineers did a much better job at keeping the runways smooth than the British teams that originally built them.

http://warbirdsnews.com/warbirds-news/poor-lambs-corsairs-baa-baa-blacksheep.html
I enjoyed watching this show as a youngster, with Corsairs flying out of what looked like  a dirt airstrip in the Channel Isands.
Filming apparently took place at Indian Dunes Airfield in CA.
Pictures of the set at the bottom of the site sure seem to show a dirt strip.
Other links about Indian Dunes Airfield discuss it being turned into farmland, I don't think that would have been done if it was paved.  Of course, if the land was worth enough, anything goes I guess.

At more sophisticated airbases, yes.  However plenty of bases used grass fields or strips as well as dirt strips plowed flat with a dozer blade.

Can you give a list of a few airbases of the USA that had unpaved runways? I haven't been able to find any.

Maybe Edwards AAF at Muroc.  There were plenty in Great Britain and Germany, though.

@CarloFelicione: Edwards still has unpaved runways (albeit also some paved runways) even today.

In 1939 most military airfields were grass strips, since there were no four-engined bombers, and most warplanes were capable of landing on roads or meadows in emergency (as instructors say, "You can land anywhere -once.")
With the outbreak of war, aircraft designers needed more powerful engines (with larger, more sensitive air intakes) and longer take-off distances (obviously concrete, being smoother than grass. works better here). Commanders were also less keen on their air support taking a day off because it was raining. So suddenly, all over Europe and other combat areas, new airfields were being built, and old ones refitted, with concrete runways; these were often called "all-weather airfields" and could be strategic objectives (this Wikipedia page refers to a unit being sent to capture the only all-weather airfield in Burma).
Though the military had construction units (the British Pioneer Corps and the US Corps of Engineers), they were often overstretched, and cililian contractors had to be used. One airfield near where I live, RAF Beaulieu, was built by local labourers. Having left school at 14, too young to be called up, they did their bit by joining construction companies for a few years, and indeed the airfield opened on time in 1941. They were, however, obviously inexperienced (both the young workers and the managers who might have never seen a concrete runway, let alone built one), and working under a lot of pressure for both time and cost. Interestingly, though the RAF squadrons who used the airfield made no complaint, the USAAF pilots found that it was not up to the standards they were used to (and it rained a lot more than back home, which also degraded performance). They did not receive much sympathy from the builders who had dodged machine-gun fire. (Some memoirs have been published, but not on the internet; a few of them are still around and telling stories in the pub).